The present invention relates to orbital operations involving Earth satellites in general, and more particularly to an improved method and apparatus for rendezvousing with and/or servicing orbital platforms and satellites, or transporting material from one orbit to another.
There are many satellites in the range of altitudes generally referred to as Low Earth Orbit (LEO), particularly proximal to the lower reaches of the Van Allen belts. One preferred band of altitudes above the Earth""s surface for LEO satellite operation is between 200 km (kilometers) and 1500 km in mid inclinations, or 200 km to 1000 km in polar inclinations.
LEO satellites may malfunction for a variety of reasons including, but not limited to, failure of booms or panels to deploy, computer or transponder failure, upper stage rocket failure, loss of orientation in relation to the sun and subsequent power loss, fundamental design flaws such as optical systems that can not focus properly or running out of fuel required for orbital station keeping or maneuvering. Currently, in most cases, a malfunctioning satellite is declared a complete loss and is replaced by a new satellite. This costs many tens of millions of dollars for commercial LEO satellites, hundreds of millions of dollars for commercial Geostationary Earth Orbit (GEO) satellites, and upwards of a billion dollars for many defense-related satellites. In addition to the cost of replacement there is also a delay caused by the need to build a replacement satellite.
In a few cases, the failure of some satellites has been remedied by in-orbit repair of the satellite or recapture of the satellite to Earth-bound repair and re-launch. In 1995, NASA used the Space Shuttle to repair faulty optics on the $1.5 billion Hubble Space Telescope. Using the Shuttle""s robotic arm to grapple the Hubble telescope in its 600 kilometer altitude orbit, astronauts put on spacesuits, went out into space and replaced major sub-components of the Hubble system. Then, in 1997, NASA used the Shuttle to perform additional in-orbit repairs on the Hubble Space Telescope to fix failed inertial navigation sub-systems and to upgrade the Hubble Space Telescope with improved optics.
The use of the Space Shuttle for a repair mission, at an estimated mission cost of $500 million, is only practical and cost-effective for satellites with an existing value of at least half a billion dollars, and then only for satellites in orbits accessible by the Shuttlexe2x80x94from about 28.6 degrees inclination to 57 degrees inclination under normal circumstances, and under 650 kilometers altitude. With the current mix and positions of satellites in orbit today, that limits this repair scenario to less than one percent of the satellites in Earth orbit.
Direct launch from the Earth of satellite servicing apparatus, using space transportation vehicles other than the Shuttle, to the orbit occupied by a malfunctioning satellite has been proposedxe2x80x94but not implementedxe2x80x94by numerous parties. Direct launch from Earth of an apparatus that can recover or service a satellite is technically feasible, but expensive. Such a servicing approach might be useful in some cases, but the cost of launching the apparatus from Earth might well be more than the replacement cost of a satellite.
One issue for operations rendezvous in orbit is minimizing the cost and time of rendezvous. One factor that complicates rendezvous is the fact that orbits xe2x80x9cprecessxe2x80x9d around the Earth (or other planetary bodies). A brief inspection of orbital mechanics shows why this is a problem. The path that an object takes in a closed-circuit orbit around a more massive body (such as a satellite around the Earth) is in the shape of an ellipse. If we consider a satellite""s orbit around the Earth, the ellipse can be defined by its semimajor axis (a) given by Equation 1,                     a        =                                            h              A                        +                          h              P                        +                          2              ⁢              R                                2                                    (                  Equ          .                      xe2x80x83                    ⁢          1                )            
and eccentricity (e) given by Equation 2,                     e        =                                            h              A                        -                          h              P                                                          h              A                        +                          h              P                        +                          2              ⁢              R                                                          (                  Equ          .                      xe2x80x83                    ⁢          2                )            
where R is the equatorial radius of the Earth, and hA and hP are the highest and lowest altitudes of the satellite above the Earth""s surface, or xe2x80x9capogeexe2x80x9d and xe2x80x9cperigeexe2x80x9d, respectively. In the special case of a circular obit, hA=hP. The location of this ellipse in space relative to the Earth can be given by its inclination (i) relative to the equatorial plane, its right ascension of the ascending node (RAAN, or xcexa9) which is measured counterclockwise in the equator plane from the direction of the vernal equinox to the point where the satellite makes its south-to-north crossing of the equatorial plane, and argument of perigee (xcfx89) which is measured in the orbit plane in the direction of the satellites motion from the ascending node to perigee. These relationships are shown in FIG. 1.
Westward precession of an orbit, taken herein to mean changes in the orbit""s RAAN, will occur over time due to perturbations caused by J2 zonal harmonics in the central attractive body""s gravitational field (xe2x80x9coblatenessxe2x80x9d). Orbits of differing altitudes, inclinations, and eccentricity will exhibit different xe2x80x9cprecession ratesxe2x80x9d (xcexa9), which for an Earth orbit can be approximated by Equation 3.                               Ω          .                =                                            -              9.9639                                                      (                                  1                  -                                      e                    2                                                  )                            2                                xc3x97                                    (                              R                                  R                  +                                                                                    h                        A                                            +                                              h                        P                                                              2                                                              )                        3.5                    ⁢          cos          ⁢                      xe2x80x83                    ⁢                      i            ⁡                          [                              degrees                                  mean                  ⁢                                      xe2x80x83                                    ⁢                  solar                  ⁢                                      xe2x80x83                                    ⁢                  day                                            ]                                                          (                  Equ          .                      xe2x80x83                    ⁢          3                )            
Orbital rendezvous, such as changing from one orbit to another of a different altitude, requires a change in velocity (xcex94V). The xcex94V that can be achieved by expending a predetermined amount of energy, e.g., burning a predetermined amount of propellant in a rocket engine, where the predetermined amount can be calculated using Equation 4,                               Δ          ⁢                      xe2x80x83                    ⁢          V                =                              g            0                    ⁢                      I            sp                    ⁢          ln          ⁢                                                    m                f                            +                              m                fuel                                                    m              f                                                          (                  Equ          .                      xe2x80x83                    ⁢          4                )            
where g0 is the Earth""s gravitational constant at sea level, ISP is the rocket engine""s specific impulse, mf is the final mass of the space vehicle, and mfuel is the mass of the fuel used in the maneuver. Since xcex94V is related to the amount propellant used, it thus affects the cost of the transfer.
Precession affects the cost of orbit transfer because changing from one orbit to another one in a different plane, i.e., one with a different inclination and/or RAAN, requires a certain xcex94V even if those orbits are otherwise identical. For example, changing a circular orbit with velocity VC and inclination i from a RAAN of xcexa91 to xcexa92 will require a xcex94V given by Equation 5,
xcex94V=2VCsin xcex8/2xe2x80x83xe2x80x83(Equ. 5)
where the equivalent plane change angle xcex8 is given by Equation 6.
cos xcex8=cos2i+sin2 cos(xcexa92xe2x88x92xcexa91)xe2x80x83xe2x80x83(Equ. 6)
By waiting for a specific orbit to precess to the same RAAN of another orbit, a minimum energy trip between the two orbits is available since no plane change is required. Conversely, reducing the waiting time needed to transfer from one orbit to another can be achieved by performing a transfer with some plane change, at the expense of extra xcex94V required.
For missions launched from Earth into LEO, the relative precession rate between the launch site and the plane of an orbit in LEO is large (since the launch site on the Earth""s surface rotates through 360 degrees in about one day) and hence the period between launch opportunities, or xe2x80x9cwindowsxe2x80x9d, that can be made given the xcex94V capability of the booster is measured in hours to a few days. For example, if a Delta rocket launch designed to emplace an Iridium satellite in a particular orbital plane is not launched on schedule, the next opportunity for launch presents itself within 24 hours. Therefore, precession rates have not presented a significant problem when traveling from Earth to LEO.
However, the relative precession rate between any two independent orbits in LEO is much smaller, and the corresponding periods between minimum energy windows is often in months or even years. For example, the relative precession rate between the MIR space station at (approximately) 400 km altitude and a Globalstar satellite at 1400 km is about 2 degrees per day with a period between minimum energy launch windows of 180 days. This has not been a problem until now because the need to move from one orbit in LEO to another orbit in LEO has been minimal. But a manned reusable space vehicle based at a space platform that was required to travel to a satellite in LEO to service or capture it will need to compromise between minimum time and cost needed to perform the mission.
Using a reusable space vehicle that is docked in an intermediate orbit for rescue missions according to one embodiment of the present invention, a satellite can be serviced with less delay, energy expenditure, and cost than a space vehicle launched from Earth for each mission. Additionally, if the reusable space vehicle is used for repairing or maintaining a plurality of satellites all orbiting at the same altitude, operating the space vehicle according to one method of the present invention, the reusable repair vehicle can be moved from one orbit to another with minimal energy expense while not having to wait for a launch window.
In one specific embodiment, the space vehicle is maintained in a docking orbit until needed. Once a servicing need is identified, a destination orbit is identified for the space vehicle and a minimum energy path is identified. If the time to the next launch window between the docking orbit and the destination orbit happens to be near enough to allow for a timely rendezvous, the space vehicle is simply moved directly to the destination orbit. However, if the next launch window is too far into the future, the space vehicle is first moved to an intermediate orbit for a dwell time and then moved to the destination orbit.
Preferably, the satellite will also be moved to an intermediate orbit during a launch window between the docking orbit and the intermediate orbit. The intermediate orbit might be specifically selected to give a particular launch window, i.e., the desired launch window can be identified and then the intermediate orbit selected from a plurality of intermediate orbits to select the orbit with the closest launch window to the desired launch window. The space vehicle remains in the intermediate orbit for a dwell time and after the dwell time, the space vehicle moves to the destination orbit.
The dwell time is preferably selected so that the launch window between the intermediate orbit and the destination orbit occurs at the end of the dwell time. Selecting the proper intermediate orbit and dwell time would allow the space vehicle to move from the docking orbit to the intermediate orbit to the destination orbit using only minimum energy launch windows (for each orbit change) in much less time than if the repair vehicle had to wait for a launch window between the docking orbit and the destination orbit.
Where the reusable space vehicle is used for xe2x80x9cmulti-hopxe2x80x9d repair missions, the destination orbit for one hop could treated as the docking orbit for the next hop, with the docking orbit for the first hop and the destination orbit following the last repair preferably both being a docking orbit normally used by the space vehicle between missions.
The space vehicle can be a vehicle designed to be piloted by humans or telerobotically. In one implementation, the inactive space vehicle is docked in an Intermediate LEO orbit (altitudes of approximately 250 km to approximately 500 km) and is used to rendezvous with objects in High LEO orbits (altitudes of approximately 500 km to approximately 1500 km) or objects in Low LEO orbits (altitudes of approximately 250 km or less).
In one embodiment of a method for arranging a mission according to the present invention, the mission architecture includes a reusable orbital transfer vehicle. One step of the method is to station such a transfer vehicle at an orbital platform, whereat human pilots can be employed to pilot the vehicle. Because humans are involved in piloting the vehicle, transfer times are typically limited to a few days, rather than the weeks, months or years available to automatically piloted vehicles. Another step is to pilot, under human control, the transfer vehicle between target satellites in high Low Earth Orbit and consumables and payloads delivered extremely low earth orbit. In some variations of the method, the transfer vehicle will visit high LEO first, and then drop down to extremely low Earth orbit. In other variations, the transfer vehicle will visit low LEO first to pick up payloads and/or consumables and then, modifying its orbital precession by time management at low LEO, the transfer vehicle will rise up to high LEO for rendezvous with a target satellite.
Other variations of the method can be performed as described herein, to form a xe2x80x9ctriangle mission architecturexe2x80x9d that allows transfer vehicle to be stored at an orbital platform between missions, with the orbital transfer vehicle returning to its xe2x80x98homexe2x80x99 platform. A triangle mission architecture, reduced mission durations allow them to be piloted by humans with less complexity. The triangle mission architecture also allows the transfer vehicle to move payloads between low LEO, high LEO and the orbital platform, wherever required, on a timely basis and allows the transfer vehicle to refuel at low altitudes where cost of fuel delivery is cheapest.
One advantage of these aspects of the present invention is that they facilitate quick, low energy transfers between orbits. One method of doing this is to manage the differences in precession rates by rendezvous targets and the space vehicle by having the space vehicle dwell in orbits with greater or lesser precession rates until the orbital plane of the next target for rendezvous, whether it is at the same inclination or at some other inclination in the GCI coordinate frame, is available for an optimum desired transfer orbit.
Another advantage is that the unique mission architecture encompassing the regression of different orbital altitudes greatly increases the cost benefit of a mission.